Most relatively small missiles in use today are propelled by solid fuel rockets as opposed to, for example, turbojet engines. The selection of a solid fuel rocket as a propulsion device has been largely dictated by two factors. First, in many instances, a turbine engine cannot be fabricated sufficiently economically as to compete with a solid fuel rocket engine. Secondly, in small size missiles, i.e., those having relatively small diameter on the order of about six inches, it is heretofore been quite difficult to manufacture an efficient turbojet engine that would fit within the six inch envelope required of the propulsion unit for such a missile.
As a consequence of the use of solid fuel rocket engines, some degree of control of the missile flight path or trajectory is lost over that which would be available were it possible to propel the missile by a gas turbine engine whose output can be readily varied. Further, even if the gas turbine engine operates relatively inefficiently, the use of such an engine greatly extends the range of the missile.
The difficulty in economically producing small diameter gas turbine engines resides not so much in the manufacture of the compressor and/or turbine section of the engine, but rather, is more apt to be attributable to the labor intensive nature of the manufacture of the combustor. Furthermore, as combustor sizes shrink to fit within some desired small envelope as the six inch envelope of a relatively small missile mentioned previously, the difficulty in achieving efficient combustion of fuel rises asymptotically. In particular, as the size or volume of a combustor is reduced, there may be insufficient volume to allow the fuel to be first vaporized completely, burned efficiently, and then mixed uniformly.
Even when the difficulties with combustor sizing can be solved, still a further difficulty presents itself, namely, the prevention of the inevitably developed hot spots that occur in such a small combustor from deleteriously affecting the turbine nozzle and shortening its life to the point of premature failure. Assuming a typical turbine inlet temperature of about 1900.degree. F., hot spot temperatures of at least 2100.degree. F. will occur.
When gasses at the hot spot temperatures contact the vanes or blades of the turbine nozzle, overheating occurs which shortens the life of the nozzle and may lead to premature malfunctioning of the turbine engine itself. Consequently, there have been a number of proposals for the cooling of turbine nozzle blades, generally by providing them with internal passages through which cooling air, typically bleed air from the compressor or air by otherwise bypassing a combustor, may flow for cooling purposes. See, for example, the disclosures of U.S. Pat. No. 4,296,599 issued Oct. 27, 1981 to Adamson and U.S. Pat. No. 4,522,557 issued June 11, 1985 to Bouillier.
These approaches to the cooling of the stationary turbine nozzle blades, while effective, are far from being economical and therefore do not lend themselves to use in turbine engines that must be economically manufactured.
The present invention is directed to overcoming one or more of the above problems.